1. Field of the Invention
The present invention relates generally to airfoils in a gas turbine engine, and more specifically to an insert located within a cooling air passage of the airfoil.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine produces mechanical work from combustion of a fuel. The gas turbine engine has a compressor to supply a compressed air to a combustor, where a fuel is mixed and burned with the compressed air to produce a hot gas flow. The hot gas flow is passed through a turbine to convert the hot gas flow into mechanical work by driving the turbine shaft.
The efficiency of the gas turbine engine can be improved by operating the turbine at higher temperatures. Because the operating temperature of the turbine is above the safe operating temperature of the materials used to make parts of the turbine, such as the blades and vanes (both considered to be airfoils), the airfoils in the turbine section are cooled by passing a fluid such as compressed air through cooling passages formed within the airfoils. Improved cooling of the airfoils can allow for higher turbine operating temperatures, resulting in improved performance.
A Prior Art turbine blade is shown in FIG. 1 with an aft flowing triple pass (3-pass) serpentine cooling passage for an all convectively cooled blade. A cross sectional view of the blade is shown in FIG. 2. In the Prior Art FIG. 1 blade 12, the blade leading edge is cooled with the first up pass of the multi-pass channel flow 14. The blade mid-chord is cooled with the second leg 15 of the serpentine down pass flow channel. The aft portion of the blade is cooled with the third leg 16 of the serpentine flow channel in conjunction with a plurality of trailing edge exit discharge cooling holes 17. As the cooling air flow rate is reduced, the internal through flow velocity within the serpentine flow channels will be reduced, resulting in a low internal heat transfer rate coefficient and low internal cooling capability. For an airfoil that is designed with large internal flow cavities and low cooling flow rate, lowering the cooling flow rate to improve efficiency would result in less cooling of the airfoil. To provide adequate cooling of the airfoil with this design, a large volume of cooling air must be passed through the airfoil. Since the cooling air for the airfoil is generally from the compressor at high pressure, much of the cooling air is wasted. One way to retain the high internal cooling performance for a low cooling flow rate design with large internal serpentine flow cavity is by reducing the internal through flow area.
It is an object of the present invention to improve the blade cooling of an airfoil that is designed for a low cooling flow rate and large internal flow cavities while still using a low cooling air flow rate.
It is another object of the present invention to provide cooling air flow to both the pressure side wall and suction side wall of the airfoil while maintaining a high flow rate through the airfoil cooling passages and therefore have a high heat transfer coefficient.